In space applications, such as rockets, satellites and other space vehicles, liquid propellant thrusters and rocket engines are often used. Such thrusters and rocket engines can for example be used for the purpose of positioning and attitude control of satellites and other space vehicles. For such purposes attitude control thrusters operating in the thrust range of typically 0.5–50 N, or Δ-V rocket engines typically operating in the thrust range of 1 N to several kN. Attitude control thrusters are required to perform short pulses or pulse trains, the duration of which typically can be fractions of seconds to several minutes.
Liquid propellants can be divided into monopropellants and bipropellants. The former consists of one component, while the latter consists of two components, i.e. a liquid oxidiser and a liquid fuel. The distinction between monopropellants a bipropellants is made according to the number of components which are injected into the engine for combustion for the specific propellant.
In the case of a monopropellant, which may be a mixture of several compounds or one single chemical, only one component is injected into the engine.
Currently, hydrazine is about the only liquid monopropellant widely used for generation of hot gases. In the case of hydrazine the decomposition pathway occurs in two stages; first hydrazine is catalytically decomposed into hydrogen and ammonia in an exothermal reaction, and thereafter ammonia further decomposes into hydrogen and nitrogen in an endothermal reaction due to the high temperature generated in the first stage. The second stage endothermal reaction will reduce the flame temperature and reduce the specific impulse. It is therefore desirable to limit the ammonia dissociation as much as possible. When the ammonia dissociation is held to 55%, the adiabatic reaction temperature will be ca. 900° C.
In a bipropellant engine, fuel and oxidizer liquids are injected, atomised and mixed in a first zone of the combustion chamber. In the case of a hypergolic bipropellant, such as hydrazine and nitrogen tetroxide, there is an initial chemical reaction in the liquid phase when a droplet of fuel impinges with a droplet of oxidiser. Bipropellants which are not hypergolic use some type of ignitor to initiate the chemical combustion. In a bipropellant system using hydrogen peroxide as the oxidiser, a catalyst may be used.
Liquid bipropellants generally offer higher specific impulse than liquid monopropellants. Bipropellant systems are thus more efficient than monopropellant systems, but tend to be more complicated because of the extra hardware components needed to make sure the proper amount of fuel is mixed with the proper amount of oxidiser.
Liquid monopropellants, based on a dinitramide compound, and especially ammonium dinitramide (ADN), have recently been developed, and are disclosed in WO0050363. These propellants are novel High Performance Monopropellants, which generate extremely high temperatures at proper combustion thereof. Such monopropellant comprises at least two components; a dinitramide compound (oxidiser) and a fuel. An additional solvent component may also be included, such as water.
These new monopropellants, including at least two components, have been described to generate a very high temperature on combustion, such as about 1700° C. As the propellant may also include water, very high demands will be put on a suitable engine or thruster for such a fuel, consequently excluding all known monopropellant thrusters as suitable alternatives.
Thus, it is an object of the present invention to provide a reactor for decomposition and combustion of liquid ammonium dinitramide-based monopropellants.
It is a further object of the present invention to provide a process for decomposition ammonium dinitramide-based monopropellants, such as for rocket propulsion and for controlled gas generation for any other purpose, such as rotary power in auxiliary power units.
Other objects and advantages of the present invention will become evident from the following description, examples, and the attached claims.
The terms rocket engine and thruster will be used interchangeably herein to designate the portion of a liquid propellant rocket engine, in which the propellant is injected, extending downstream to the nozzle.